Gas turbine power plants are used as the primary propulsive power source for aircraft, in the forms of jet engines and turboprop engines, as auxiliary power sources for driving air compressors, hydraulic pumps, etc. on aircraft, and as stationary power supplies such as backup electrical generators for hospitals and the like. The same basic power generation principles apply for all of these types of gas turbine power plants. A gas turbine engine in its basic form includes a compressor section, a combustion section and a turbine section arranged to provide a generally axially extending flow path for the working gases. Compressed air is mixed with fuel and burned, and the energetic hot combustion gases are directed against stationary turbine guide vanes in the one or more turbine stages of the engine. The vanes turn the high velocity gas flow partially sideways to impinge at the proper angle upon turbine blades mounted on a turbine disk or wheel that is free to rotate.
The force of the impinging gas causes the turbine disk to spin at high speed. The power so generated is then used to draw more air into the engine, in the case of the jet propulsion engine, and both draw more air into the engine and also supply shaft power to turn the propeller, an electric generator, or for other uses, in the cases of the other applications. The high velocity combustion gas is then passed out the aft end of the gas turbine which, in the propulsion engine applications, also supplies a forward reaction force to the aircraft.
As is well known, the thermal efficiency, and therefore power, produced by any engine is a function of, among other parameters, the temperature of the working gases admitted into the turbine section. That is, all other things being equal, an increase in power from a given engine can be obtained by increasing the combustion gas temperature. This is particularly true for small turboshaft or turboprop engines where very small changes in the operating temperature can substantially affect the engine output. For example, it has been determined in a typical engine of this type that a single degree centigrade increase in the temperature of the working gases can increase the engine power by as much as 15 horsepower. However, as a practical matter, the maximum feasible gas temperature, and hence the efficiency and output of the engine, is limited by the high temperature capabilities of the various turbine section components exposed to the hot gas flow.
The turbine blades and vanes lie at the heart of the power plant, and it is well established that in most cases they are the limiting factors in achieving improved power plant efficiency. In particular, because they are subjected to high heat and stress loadings as they are impacted by the hot gas, there is a continuing effort to identify improvements to the construction and materials of turbine blades and vanes to achieve ever higher performance.
In order to achieve more of the power theoretically available from higher turbine temperatures, the blades and vanes must be effectively cooled to a safe temperature which is considerably less than the maximum working gas temperature. The prior art has provided many designs for supplying a flow of air, typically bled from the compressor section, to cooling passages within the blades and/or vanes. See, for example, U.S. Pat. Nos. 4,136,516: 4,162,136: 4,505,640: 4,684,322: 4,759,688 and 4,786,234.
However, any use of the compressor air for cooling decreases the amount available for combustion, reducing the maximum power available. Furthermore, any work or heat added to this air is energy lost to the turbine cycle, thus reducing overall thermal efficiency. It would therefore be desirable to adjust the amount of air drawn off for cooling depending on the need for such: that is in response to variations in engine operating conditions. Accordingly, attempts have been made to achieve some degree of control of the cooling air flows by making use of the known dimensional variations of certain elements of the cooling system when they are subjected to the variations in engine speed or, more directly, gas temperature. See, for example, U.S. Pat. Nos. 3,712,756: 4,213,738 and 4,730,982.
All of these designs possess significant problems in making efficient use of the cooling air in that they require a preliminary design choice to be made concerning the amount of cooling necessary for each group or type of component. That is, the maximum exposure temperature during the severest engine operating conditions, such as during take-off when engine speed and loading is highest, is determined and compared to the permissible operating temperature of the components. The calculated amount of cooling air necessary to achieve adequate cooling under those transient conditions is provided and usually increased even more to provide a margin of safety and/or a lengthened operating life for the components. Once an engine is designed with the desired cooling flows established, it is difficult and/or expensive to change even if full scale engine tests determine that a different amount of cooling would be optimum.
Another problem with the prior art designs is that they assume a known or uniform temperature distribution around the central axis of the turbine. However, as will be explained later, such an assumption is generally not accurate. Circumferential variations in gas temperature are not as much a problem for the rapidly rotating turbine blades, which are affected by the average gas temperature, but seriously affect the cooling needs and/or the life of the stationary guide vanes.
Thus it is an object of the present invention to provide a method and apparatus for adjusting the flow of cooling air to a turbine component depending on the need therefore so as to conserve cooling air.
Another object of the invention is to provide means for varying the cooling air flow to turbine components in response to the temperature of the components so as to avoid overheating.
A further object of the invention is to provide a means for supplying a varying amount of cooling air to each of a series of turbine guide vanes to compensate for circumferential variations in combustion gas temperature.